Gas turbine engine airfoil

ABSTRACT

An airfoil arrangement of a turbine engine according to an example of the present disclosure includes adjacent airfoils including pressure and suction sides extending in a radial direction from a 0% span position to a 100% span position. The airfoils have a relationship between a gap/chord ratio and span position that defines a curve with a gap/chord ratio having a portion with a negative slope.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.14/626,130, filed Feb. 19, 2015, which claims the benefit of U.S.Provisional Application No. 61/941,671, which was filed on Feb. 19, 2014and is incorporated herein by reference.

BACKGROUND

This disclosure relates to gas turbine engine airfoils. Moreparticularly the disclosure relates to airfoil gap/chord ratio in, forexample, a gas turbine engine compressor or fan section.

A turbine engine such as a gas turbine engine typically includes a fansection, a compressor section, a combustor section and a turbinesection. Air entering the compressor section is compressed and deliveredinto the combustor section where it is mixed with fuel and ignited togenerate a high-speed exhaust gas flow. The high-speed exhaust gas flowexpands through the turbine section to drive the compressor and the fansection. The compressor section typically includes at least low and highpressure compressors, and the turbine section includes at least low andhigh pressure turbines.

Direct drive gas turbine engines include a fan section that is drivendirectly by one of the turbine shafts. Rotor blades in the fan sectionand a low pressure compressor of the compressor section of direct driveengines rotate in the same direction.

Gas turbine engines have been proposed in which a geared architecture isarranged between the fan section and at least some turbines in theturbine section. The geared architecture enables the associatedcompressor of the compressor section to be driven at much higherrotational speeds, improving overall efficiency of the engine. Thepropulsive efficiency of a gas turbine engine depends on many differentfactors, such as the design of the engine and the resulting performancedebits on the fan that propels the engine and the compressor sectiondownstream from the fan. Physical interaction between the fan and theair causes downstream turbulence and further losses. Although some basicprinciples behind such losses are understood, identifying and changingappropriate design factors to reduce such losses for a given enginearchitecture has proven to be a complex and elusive task.

Similarly, the fan section can also be a significant noise source, asnoise is produced by fluid dynamic interaction between the fan bladesand the incoming air stream. Some fan blade arrangements have channelsthat converge at a location downstream of the fan blade leading edgesfor most or all span positions in an attempt to reduce noise. However,fan blade arrangements that may attempt to mitigate noise may come atthe expense of reduced propulsive efficiency.

Prior compressor airfoil geometries may not be suitable for thecompressor section of gas turbine engines using a geared architecture,since the significantly different speeds of the compressor changes thedesired aerodynamics of the airfoils within the compressor section.Counter-rotating fan and compressor blades, which may be used in gearedarchitecture engines, also present design challenges.

SUMMARY

An airfoil arrangement of a turbine engine according to an example ofthe present disclosure includes adjacent airfoils including pressure andsuction sides extending in a radial direction from a 0% span position toa 100% span position. The airfoils have a relationship between agap/chord ratio and span position that defines a curve with a gap/chordratio having a portion with a negative slope. The adjacent airfoils arerotatable blades.

In a further embodiment of any of the foregoing embodiments, thegap/chord ratio is less than 1.0 at 100% span.

In a further embodiment of any of the foregoing embodiments, thegap/chord ratio is less than 1.0 at each span position.

In a further embodiment of any of the foregoing embodiments, thegap/chord ratio is less than 0.9 at 0% span.

In a further embodiment of any of the foregoing embodiments, thegap/chord ratio is less than 1.04 from 80% span to 100% span.

In a further embodiment of any of the foregoing embodiments, thegap/chord ratio is less than 0.7 at 0% span.

In a further embodiment of any of the foregoing embodiments, the portionis from 80% span to 100% span.

A compressor section according to an example of the present disclosureincludes a plurality of rotatable blades and a plurality of vanesarranged about an axis, wherein adjacent blades include pressure andsuction sides extending in a radial direction from a 0% span position toa 100% span position. The adjacent blades have a relationship between agap/chord ratio and span position that defines a curve with a gap/chordratio having a portion with a negative slope.

In a further embodiment of any of the foregoing embodiments, thegap/chord ratio is less than 1.0 at 100% span.

In a further embodiment of any of the foregoing embodiments, thegap/chord ratio is less than 0.7 at 0% span.

In a further embodiment of any of the foregoing embodiments, thegap/chord ratio is less than 1.0 at each span position.

In a further embodiment of any of the foregoing embodiments, thegap/chord ratio is less than 1.04 from 80% span to 100% span, and theportion is from 80% span to 100% span.

A gas turbine engine according to an example of the present disclosureincludes a combustor section arranged between a compressor section and aturbine section, a fan section including an array of fan blades, andadjacent airfoils including pressure and suction sides extending in aradial direction from a 0% span position to a 100% span position. Theairfoils have a relationship between a gap/chord ratio and span positionthat defines a curve with a gap/chord ratio having a portion with anegative slope, and the gap/chord ratio is less than 1.04 at 100% span.

In a further embodiment of any of the foregoing embodiments, the fansection includes twenty-six or fewer fan blades, and has a fan pressureratio that is less than 1.55.

In a further embodiment of any of the foregoing embodiments, theairfoils are rotatable relative to an engine static structure.

In a further embodiment of any of the foregoing embodiments, theairfoils are arranged in the compressor section.

In a further embodiment of any of the foregoing embodiments, thegap/chord ratio is greater than 0.6 at each span position.

In a further embodiment of any of the foregoing embodiments, thecompressor section includes a first compressor section and a secondcompressor section. The second compressor section is arrangedimmediately upstream of the combustor section, and the airfoils areprovided in the first compressor section.

In a further embodiment of any of the foregoing embodiments, the firstcompressor section is immediately downstream from the fan section and iscounter-rotating relative to the array of fan blades.

A further embodiment of any of the foregoing embodiments includes a geararrangement configured to drive the fan section, wherein the turbinesection is configured to drive the gear arrangement and the compressorsection.

In a further embodiment of any of the foregoing embodiments, thegap/chord ratio is less than 0.9 at 0% span, and the gap/chord ratio isless than 1.04 from 80% span to 100% span.

In a further embodiment of any of the foregoing embodiments, the portionis from 80% span to 100% span, and the curve has another portion with apositive slope.

These and other features of this disclosure will be better understoodupon reading the following specification and drawings, the following ofwhich is a brief description.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure can be further understood by reference to the followingdetailed description when considered in connection with the accompanyingdrawings wherein:

FIG. 1 schematically illustrates a gas turbine engine embodiment with ageared architecture.

FIG. 2 schematically illustrates a low pressure compressor section ofthe gas turbine engine of FIG. 1.

FIG. 3 is a schematic view of airfoil span positions.

FIG. 4A is a schematic view of a cross-section of an airfoil sectionedat a particular span position and depicting directional indicators.

FIG. 4B is a cross-sectional view of a sheath arrangement for anairfoil.

FIG. 5 is a schematic view of adjacent airfoils depicting a gap and achord of the airfoil.

FIG. 6A is a perspective view of a fan section.

FIG. 6B is a schematic cross-sectional view of the fan section of FIG.6A.

FIG. 7 graphically depicts curves for several example airfoil gap/chordratios relative to span, including two prior art curves and severalinventive curves according to this disclosure.

FIG. 8A is a schematic view of the adjacent airfoils depicting pointpairs along a channel between the adjacent airfoils at a first spanposition.

FIG. 8B graphically depicts curves for example airfoil chord length topoint pair ratios relative to engine position between the adjacentairfoils of FIG. 8A.

FIG. 8C is a schematic view of the adjacent airfoils depicting pointpairs along a channel between the adjacent airfoils at a second spanposition.

FIG. 8D graphically depicts curves for example airfoil chord length topoint pair ratios relative to engine position between the adjacentairfoils of FIG. 8C.

FIG. 8E is a schematic view of the adjacent airfoils depicting pointpairs along a channel between the adjacent airfoils at a third spanposition.

FIG. 8F graphically depicts curves for example airfoil chord length topoint pair ratios relative to engine position between the adjacentairfoils of FIG. 8E.

The embodiments, examples and alternatives of the preceding paragraphs,the claims, or the following description and drawings, including any oftheir various aspects or respective individual features, may be takenindependently or in any combination. Features described in connectionwith one embodiment are applicable to all embodiments, unless suchfeatures are incompatible.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmenter section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures. That is, the disclosedairfoils may be used for engine configurations such as, for example,direct fan drives, or two- or three-spool engines with a speed changemechanism coupling the fan with a compressor or a turbine sections.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis X relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisX which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five (5:1). Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicyclic geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

The example gas turbine engine includes the fan 42 that comprises in onenon-limiting embodiment less than about twenty-six (26) fan blades. Inanother non-limiting embodiment, the fan section 22 includes less thanabout twenty (20) fan blades. Moreover, in one disclosed embodiment thelow pressure turbine 46 includes no more than about six (6) turbinerotors schematically indicated at 34. In another non-limiting exampleembodiment the low pressure turbine 46 includes about three (3) turbinerotors. A ratio between the number of fan blades 42 and the number oflow pressure turbine rotors is between about 3.3 and about 8.6. Theexample low pressure turbine 46 provides the driving power to rotate thefan section 22 and therefore the relationship between the number ofturbine rotors 34 in the low pressure turbine 46 and the number ofblades 42 in the fan section 22 disclose an example gas turbine engine20 with increased power transfer efficiency.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.55. In another non-limiting embodimentthe low fan pressure ratio is less than about 1.45. In anothernon-limiting embodiment the low fan pressure ratio is from 1.1 to 1.45.“Low corrected fan tip speed” is the actual fan tip speed in ft/secdivided by an industry standard temperature correction of [(Tram °R)/(518.7° R)]^(0.5). The “Low corrected fan tip speed” as disclosedherein according to one non-limiting embodiment is less than about 1200ft/second (365.7 meters/second).

Referring to FIG. 2, which schematically illustrates an example lowpressure compressor (LPC) 44, a variable inlet guide vane (IGV) isarranged downstream from a fan exit stator (FES). The figure is highlyschematic, and the geometry and orientation of various features may beother than shown. An actuator driven by a controller actuates the IGVabout their respective axes. Multiple airfoils are arranged downstreamfrom the IGV. The airfoils include alternating stages of rotors (ROTOR1,ROTOR2, ROTOR3, ROTOR4) and stators (STATOR1, STATOR2, STATOR3,STATOR4). In the example shown in FIG. 2, the LPC includes four rotorsalternating with four stators. However, in another example, a differentnumber of rotors and a different number of stators may be used.Moreover, the IGV and stator stages may all be variable, fixed or acombination thereof.

The disclosed airfoils may be used in a low pressure compressor of a twospool engine or in portions of other compressor configurations, such aslow, intermediate and/or high pressure areas of a three spool engine.However, it should be understood that any of the disclosed airfoils maybe used for blades or vanes, and in any of the compressor section,turbine section and fan section.

In some examples, the fan section 22 includes a hardwall containmentsystem 23 arranged about the engine axis A and spaced radially from thefan blades 42. The hardwall containment system 23 is configured tocontain, and absorb the impact of, a fan blade 42 separating from a fanhub 77 or a fragment thereof. In some embodiments, the hardwallcontainment system 23 is a hard ballistic liner applied to the nacelleor fan case 15. The hard ballistic liner can include a rigid materialsuch as a resin impregnated fiber structure, metallic structures, orceramic structures.

Various materials and structures of the fan case 15 and/or hardwallcontainment system 23 can be utilized. In some examples, the fan section22 includes a composite fan case 15 made of an organic matrix composite.The organic matrix composite can include a matrix material andreinforcement fibers distributed through the matrix material. Thereinforcement fibers may be discontinuous or continuous, depending uponthe desired properties of the organic matrix composite, for example. Thematrix material may be a thermoset polymer or a thermoplastic polymer.The reinforcement fibers may include carbon graphite, silica glass,silicon carbide, or ceramic. Given this description, one of ordinaryskill in the art will recognize that other types of matrix materials andreinforcement fibers may be used. The disclosed arrangements of thecomposite fan case 15 reduce the overall weight of the nacelle assembly,thereby improving propulsive efficiency and overall performance.

Referring to FIG. 3, span positions on an airfoil 64 are schematicallyillustrated from 0% to 100% in 10% increments. The airfoil 64 can belocated in the fan section 22 or the compressor section 24, for example.Each section at a given span position is provided by a conical cut thatcorresponds to the shape of the core flow path, as shown by the largedashed lines. In the case of an airfoil with an integral platform, the0% span position corresponds to the radially innermost location wherethe airfoil meets the fillet joining the airfoil to the inner platform.In the case of an airfoil without an integral platform, the 0% spanposition corresponds to the radially innermost location where thediscrete platform meets the exterior surface of the airfoil. Forairfoils having no outer platform, such as blades, the 100% spanposition corresponds to the tip 66. For airfoils having no platform atthe inner diameter, such as cantilevered stators, the 0% span positioncorresponds to the inner diameter location of the airfoil. For stators,the 100% span position corresponds to the outermost location where theairfoil meets the fillet joining the airfoil to the outer platform.

Airfoils in each stage of the fan section or LPC are specificallydesigned radially from an inner airfoil location (0% span) to an outerairfoil location (100% span) and along circumferentially oppositepressure and suction sides 72, 74 extending in chord between a leadingand trailing edges 68, 70 (see FIG. 4A). Each airfoil is specificallytwisted with a corresponding stagger angle and bent with specific sweepand/or dihedral angles along the airfoil. Airfoil geometric shapes,stacking offsets, chord profiles, stagger angles, sweep and dihedralangles, among other associated features, are incorporated individuallyor collectively to improve characteristics such as aerodynamicefficiency, structural integrity, and vibration mitigation, for example,in a gas turbine engine with a geared architecture in view of the higherLPC rotational speeds or lower fan rotational speeds.

The airfoil 64 has an exterior surface 76 providing a contour thatextends from a leading edge 68 generally aftward in a chord-wisedirection H to a trailing edge 70, as shown in FIG. 4. Pressure andsuction sides 72, 74 join one another at the leading and trailing edges68, 70 and are spaced apart from one another in an airfoil thicknessdirection T. An array of airfoils 64 are positioned about the axis X(corresponding to an X direction) in a circumferential or tangentialdirection Y. Any suitable number of airfoils may be used for aparticular stage in a given engine application.

The exterior surface 76 of the airfoil 64 generates lift based upon itsgeometry and directs flow along the core flow path C. The airfoil 64 maybe constructed from a composite material, or an aluminum alloy ortitanium alloy, or a combination of one or more of these.Abrasion-resistant coatings or other protective coatings may be appliedto the airfoil. The rotor stages may constructed as an integrally bladedrotor, if desired, or discrete blades having roots secured withincorresponding rotor slots of a disc. The stators may be provided byindividual vanes, clusters of vanes, or a full ring of vanes.

FIG. 4B illustrates a schematic cross-sectional view of a compositeairfoil 64′ which can be utilized in the fan section 22, for example. Insome examples, the airfoil 64′ is made of a two dimensional orthree-dimensional composite. The composite may be formed from aplurality of braided yarns such as carbon fibers. Other materials can beutilized, such as fiberglass, Kevlar®, a ceramic such as Nextel™, and apolyethylene such as Spectra®. In other examples, the composite isformed from a plurality of uni-tape plies or a fabric. The fabric caninclude woven or interlaced fibers, for example.

In some examples, the airfoil 64′ is three-dimensional composite free ofa central core. In other examples, the airfoil 64′ includes one or morecores 75. The core 75 can include a foam or other lightweight materialsuch as polyurethane. Other materials can be utilized, such as metallicfoam and polymethacrylimide (PMI) foam sold under the trade nameRohacell®. In other examples, the core 75 is formed from one or moreplies of fabric or from braided yarns.

Each airfoil 64′ can include a sheath 95. In some examples, a sheath 95a is located at a leading edge 68 of the airfoil 64′. In other examples,a sheath 95 b is located at a trailing edge 70 of the airfoil 64′. Otherlocations of the sheath 95 are contemplated, such as on pressure and/orsuction sides 72, 74 of the airfoil 64′. In another example, a sheath 95extends across at least a portion of the airfoil tip 66 between theleading edge 68 and trailing edge 70 of the airfoil 64′. Variousmaterials of the sheath 95 can be utilized, such as titanium, a steelalloy or another material.

Airfoil geometries can be described with respect to various parametersprovided. The disclosed graph(s) illustrate the relationships betweenthe referenced parameters within 10% of the desired values, whichcorrespond to a hot aerodynamic design point for the airfoil. In anotherexample, the referenced parameters are within 5% of the desired values,and in another example, the reference parameters are within 2% of thedesired values. It should be understood that the airfoils may beoriented differently than depicted, depending on the rotationaldirection of the blades. The signs (positive or negative) used, if any,in the graphs of this disclosure are controlling and the drawings shouldthen be understood as a schematic representation of one example airfoilif inconsistent with the graphs. The signs in this disclosure, includingany graphs, comply with the “right hand rule.”

FIG. 5 shows an isolated view of a pair of adjacent airfoils 64. Asshown, the airfoil 64 is sectioned at a radial position between the rootand the tip. A chord 80 is shown on the section of the airfoil 64. Thechord 80, which is the length between the leading and trailing edges 68,70, forms an angle, or stagger angle α, with a tangential plane (in theY-direction) normal to the engine's central longitudinal axis in theX-direction. A dimension of the chord 80 may vary along the span of theairfoil 64. The leading edges 68 of the adjacent airfoils 64 areseparated by a gap 82 or circumferential pitch in the Y-direction. Thegap 82 is equivalent to the arc distance between the leading edges 68 ofneighboring airfoils 64 for a corresponding span position. A ratio ofgap/chord, the inverse of which is referred to as solidity, varies withposition along the span, and varies between a hot, running condition anda cold, static (“on the bench”) condition.

FIGS. 6A-6B shows an example in which the airfoils 64 are a plurality offan blades in the fan section 22. The fan 42 includes a rotor 69 havingan array or row 71 of fan blades or airfoils 64 that extendcircumferentially around and are supported by the fan hub 77. Anysuitable number of airfoils 64 may be used in a given application. Thehub 77 is rotatable about the engine axis A. The array 71 of airfoils 64are positioned about the axis A in a circumferential or tangentialdirection Y. Each of the airfoils 64 includes an airfoil body thatextends in a radial span direction R from the hub 77. A root 78 of theairfoil 64 is received in a correspondingly shaped slot in the hub 77.The airfoil 64 extends radially outward of a platform 79, which providesthe inner flow path. The platform 79 may be integral with the airfoil 64or separately secured to the hub 77, for example. A spinner 85 issupported relative to the hub 77 to provide an aerodynamic inner flowpath into the fan section 22.

The geared architecture 48 of the disclosed example permits the fan 42to be driven by the low pressure turbine 46 through the low speed spool30 at a lower angular speed than the low pressure turbine 46, whichenables the LPC 44 to rotate at higher, more useful speeds. Thegap/chord ratio in a hot, running condition along the span of theairfoils 64 provides necessary fan or compressor operation in cruise athigher speeds enabled by the geared architecture 48, to enhanceaerodynamic functionality and thermal efficiency. As used herein, thehot, running condition is the condition during cruise of the gas turbineengine 20. For example, the gap/chord ratio in the hot, runningcondition can be determined in a known manner using finite elementanalysis.

FIG. 7 illustrates the relationship between the gap/chord ratio and theaverage span (AVERAGE SPAN %), which is the average of the radialposition at the leading and trailing edges 68, 70. In one example, theairfoils are LPC rotor blades. In alternative examples, the airfoils arefan blades. Two prior art curves (“PRIOR ART”) are illustrated as wellas several example inventive curves 88, 90, 92, 94, 96. The airfoil 64has a relationship between a gap/chord ratio and span position. Thecurve has a gap/chord ratio with a portion having a negative slope,unlike the prior art entirely positive slopes. In one example, theportion is from 80% span to 100% span. The gap/chord ratio of less than1.04 from 80% span to 100% span, which is significantly less than theprior art. In one example, the gap/chord ratio that is less than 1.0 at100% span.

The initial gap/chord ratio is also less than in prior art airfoils. Inone example, the gap/chord ratio of less than 0.9 at 0% span, and inanother example, the gap/chord ratio of less than 0.8 at 0% span. Inanother example, the gap/chord ratio of less than 0.7 at 0% span.

In examples, in which the airfoil is located in the compressor section,the gap/chord ratio of the inventive curves is less than prior artairfoils due to the higher speed of the LPC, which maintains theefficiency of the airfoils. A larger gap/chord ratio than the inventivecurves, such as those similar to prior art airfoils, would permit thecore flow to pass through the blades without sufficiently compressingthe fluid, thereby reducing compressor effectiveness.

FIGS. 8A-8E illustrate example channel arrangements for an array ofairfoils. Referring to FIG. 8A, a leading airfoil 64A and a followingairfoil 64B are spaced apart in the circumferential direction Y todefine a channel 96 extending generally in a chordwise direction from aleading edge 68 of airfoil 64B. Pressure side 72 of airfoil 64B andsuction side 74 side of airfoil 64A define a plurality of segments orchannel widths 98 along the channel 96. For the purposes of thisdisclosure, each channel width 98 along the channel 96 is astraight-line connected point pair extending from a point 101B on thepressure side 72 of the airfoil 64B to a point 101A on the suction side74 of the airfoil 64A that is a minimum distance from the point on thepressure side 72. In this manner, a connected point pair can be definedat each location along the channel 96. The channel widths 98 varygenerally in the axial direction X due to contouring of exteriorsurfaces 76. The channel widths 98 may vary along the span of theairfoil 64.

The channel 96 is provided with an inlet 99 at the leading edge 68 ofairfoil 64B and an outlet 100 downstream of the inlet 99. In someexamples, the width of the channel 96 diverges without converging in achordwise direction along the channel 96 for at least some of the spanpositions. In further examples, the width of the channel 96 divergeswithout converging in a chordwise direction along the channel 96 foreach of the span positions. This arrangement is shown in FIG. 8E at agiven span position. Rather, a minimum width of the channel 96 from theinlet 99 to the outlet 100 of the channel 96 increases from 0% span to100% span.

In other examples, at some span positions the channel width 98 convergesalong the channel 96 to define a nozzle 102, where 102 is labeled at athroat of the nozzle, configured to meter flow of the incoming airthrough the channel 96. The nozzle 102 defines a minimum channel width98 along the channel 96. The nozzle 102 is located downstream of aposition along the pressure side 72 of airfoil 64B at a radius definedby the leading edge 68. Rather, the inlet 99 is characterized in part bythe geometry of the leading edge 68 (best seen in FIG. 8D), whereas thenozzle 102 is characterized by the contouring of the pressure side 72downstream of the inlet 99 (best seen in FIG. 8A). In some examples, thechannel width 98 converges and diverges along the nozzle 102, commonlyreferred to as a “venturi” or “converging-diverging” configuration, forat least some span positions (shown in FIG. 8A). In further examples,the channel width 98 at the nozzle 102 increases as span positionincreases. In other examples, the channel 96 has substantially the sameminimum channel width 98 from the nozzle 102 to the outlet 100, commonlyreferred to as a “converging-straight” configuration.

Various arrangements for the nozzle 102 can be utilized in accordancewith the teachings herein. In one example, the nozzle 102 extendsradially from about a 0% span position. In another example, the nozzle102 is spaced radially from the 0% span position. In some examples, thechannel width 98 converges and diverges along the channel 96 at spanpositions greater than 5% span, greater than 10% span, or greater thanabout 20% span. In another example, the nozzle 102 extends, less than orequal to about 50% of the span positions. In yet another example, thenozzle 102 extends less than or equal to about 20% of the spanpositions, and in further examples, the nozzle 102 extends radiallyoutward less than or equal to about 20% span. In some examples, thechannel width 98 diverges without converging for greater than or equalto about 80% of the span positions. In some examples, the channel width98 diverges without converging at the tip 66 or 100% span such that anozzle is not formed at the tip 66. In further examples, the channelwidth 98 diverges without converging from 100% span to about 80% span,or from 100% span to about 50% span.

FIG. 8B illustrates example plots 103A, 103B for a ratio between channelwidths (O) and a dimension (tau) of the gap 82 or circumferential pitchcorresponding to airfoils 64A, 64B of FIG. 8A at 0% span position. Thechannel widths (O) correspond to positions along the along the channel96, including locations 98, 99, 100 and 102. As shown, the ratio (O/tau)decreases from the inlet 99 to nozzle 102 with respect to the engineaxis A (x-axis) and thereafter increases downstream of the nozzle 102 tothe outlet 100 to define inflections 105A, 105B. Rather, the ratio(O/tau) at the nozzle 102 is the minimum value for the channel widths 98along the channel 96. The nozzle 102 is defined for less than half ofthe span of the channel 96 such that the channel width 98 increases oris divergent from the inlet 99 to the outlet 100 for other portions ofthe span such that the ratio (O/tau) generally increases from the inlet99 to the outlet 100 of the channel 96, as illustrated by FIGS. 8C-8D atabout 20% span, for example. Other portions of the channel 96 aredivergent or increase from the inlet 99 to the outlet 100 of thechannel, as illustrated by FIGS. 8E-8F at about 100% span, according toan example.

Engines made with the disclosed architecture, and including thecompressor and fan section arrangements as set forth in thisapplication, and with modifications coming from the scope of the claimsin this application, thus provide very high efficient operation, havereduced noise emissions, and are compact and lightweight relative totheir thrust capability. Two-spool and three-spool direct drive enginearchitectures can also benefit from the teachings herein.

It should also be understood that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom. Although particular step sequencesare shown, described, and claimed, it should be understood that stepsmay be performed in any order, separated or combined unless otherwiseindicated and will still benefit from the present invention.

Although the different examples have specific components shown in theillustrations, embodiments of this invention are not limited to thoseparticular combinations. It is possible to use some of the components orfeatures from one of the examples in combination with features orcomponents from another one of the examples.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of the claims. For that reason, the following claimsshould be studied to determine their true scope and content.

What is claimed is:
 1. An airfoil arrangement of a turbine enginecomprising: adjacent airfoils including pressure and suction sidesextending in a radial direction from a 0% span position to a 100% spanposition, wherein the airfoils have a relationship between a gap/chordratio and span position that defines a curve with a gap/chord ratiohaving a portion with a negative slope, wherein the gap/chord ratio isless than 0.8 from 80% span to 100% span, and the adjacent airfoils arerotatable blades.
 2. The airfoil arrangement according to claim 1,wherein the gap/chord ratio is less than 1.0 at each span position. 3.The airfoil arrangement according to claim 1, wherein the gap/chordratio is less than 0.9 at 0% span.
 4. The airfoil arrangement accordingto claim 1, wherein the gap/chord ratio is less than 0.7 at 0% span. 5.The airfoil arrangement according to claim 1, wherein the portion isfrom 80% span to 100% span.
 6. A compressor section comprising: aplurality of rotatable compressor blades and a plurality of compressorvanes arranged about an axis; wherein adjacent rotatable compressorblades include pressure and suction sides extending in a radialdirection from a 0% span position to a 100% span position, wherein theadjacent rotatable compressor blades have a relationship between agap/chord ratio and span position that defines a curve with a gap/chordratio having a portion with a negative slope, wherein the gap/chordratio is less than 1.04 from 80% span to 100% span.
 7. The compressorsection according to claim 6, wherein the gap/chord ratio is less than0.7 at 0% span.
 8. The compressor section according to claim 7, whereinthe gap/chord ratio is less than 1.0 at each span position.
 9. Thecompressor section according to claim 6, wherein the portion is from 80%span to 100% span.
 10. The compressor section according to claim 9,wherein the portion is from 80% span to 100% span, the curve has anotherportion with a positive slope, and the gap/chord ratio is less than 0.8at 100% span.
 11. A gas turbine engine comprising: a combustor sectionarranged between a compressor section and a turbine section; a fansection including an array of fan blades; and said compressor sectionincluding rotatable blades defining adjacent airfoils, the adjacentairfoils including pressure and suction sides extending in a radialdirection from a 0% span position to a 100% span position, wherein theairfoils have a relationship between a gap/chord ratio and span positionthat defines a curve with a gap/chord ratio having a portion with anegative slope, and the gap/chord ratio is less than 1.04 from 80% spanto 100% span, and the adjacent airfoils are rotatable blades.
 12. Thegas turbine engine according to claim 11, wherein the fan sectionincludes twenty-six or fewer fan blades, and has a fan pressure ratiothat is less than 1.55.
 13. The gas turbine engine according to claim11, wherein the airfoils are rotatable relative to an engine staticstructure.
 14. The gas turbine engine according to claim 11, wherein thegap/chord ratio is greater than 0.6 at each span position.
 15. The gasturbine engine according to claim 11, wherein the compressor sectionincludes a first compressor section and a second compressor section, thesecond compressor section arranged immediately upstream of the combustorsection, and the airfoils are provided in the first compressor section.16. The gas turbine engine according to claim 15, wherein the firstcompressor section is immediately downstream from the fan section and iscounter-rotating relative to the array of fan blades.
 17. The gasturbine engine according to claim 15, wherein the gap/chord ratio isless than 0.8 at 100% span.
 18. The gas turbine engine according toclaim 17, wherein the portion is from 80% span to 100% span, the curvehas another portion with a positive slope, and the gap/chord ratio isless than 0.8 at 0% span.
 19. The gas turbine engine according to claim11, comprising a gear arrangement configured to drive the fan section,wherein the turbine section is configured to drive the gear arrangementand the compressor section.
 20. The gas turbine engine according toclaim 11, wherein the gap/chord ratio is less than 0.9 at 0% span. 21.The gas turbine engine according to claim 11, wherein the portion isfrom 80% span to 100% span, and the curve has another portion with apositive slope.